Gas turbine engine with active variable turbine cooling

ABSTRACT

A gas turbine engine includes a compressor section, a combustor section, and a turbine section operably coupled to the compressor section. A primary flow path is defined through the compressor section, the combustor section, and the turbine section. An engine case surrounds the compressor section, the combustor section, and the turbine section. The gas turbine engine also includes a means for providing an active variable cooling flow through a bypass duct external to the engine case to a secondary flow cavity of the turbine section.

BACKGROUND

Exemplary embodiments pertain to aircraft systems, and more particularlyto systems and methods of actively cooling a turbine section of a gasturbine engine.

Gas turbine engines can provide propulsion and power on an aircraft. Gasturbine engines can be implemented as Brayton cycle machines withbalanced thermodynamic cycles, where work is a function of pressure andvolume, and heat transfer is balanced. The net work for a thermodynamicexchange in a gas turbine engine may be expressed as work done on asubstance due to expansion minus work done on recompression. Work can belost at thermodynamic exchanges where a cooling air branch occurswithout imparting a motive force to turbomachinery within a gas turbineengine. In some instances, pressures and temperatures within a gasturbine engine are constrained due to material properties, which canresult in designs that are less efficient through losses than mayotherwise be needed. There are parts in high temperature regions in agas turbine engine that require active cooling, which is detrimental toperformance and fuel efficiency. Typically, these high-temperaturecooled parts are designed under worst case, highest temperatureconditions.

BRIEF DESCRIPTION

Disclosed is a gas turbine engine that includes a compressor section, acombustor section, and a turbine section operably coupled to thecompressor section. A primary flow path is defined through thecompressor section, the combustor section, and the turbine section. Anengine case surrounds the compressor section, the combustor section, andthe turbine section. The gas turbine engine also includes a means forproviding an active variable cooling flow through a bypass duct externalto the engine case to a secondary flow cavity of the turbine section.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the means forproviding the active variable cooling flow includes a cooling airmetering valve.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the cooling airmetering valve is electronically actuated based on either or both of aflight phase and an operating parameter of the gas turbine engine.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the means forproviding the active variable cooling flow includes an airflow pathwithin a static structure of the gas turbine engine between the bypassduct and the secondary flow cavity of the turbine section.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the airflowpath includes an inner diffuser flow path configured to deliver ametered supply of cooling air from the compressor section through adiffuser section proximate to the combustor section.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the secondaryflow cavity of the turbine section includes a static support structurecooling supply cavity of at least one stage of vanes of the turbinesection.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the means forproviding the active variable cooling flow includes a compressor bleedport in the engine case proximate to the compressor section.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the bypass ductextends external to the engine case around the combustor section fromthe compressor bleed port to at least one static support structurecooling supply cavity of the turbine section.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the activevariable cooling flow is extracted from a diffuser section proximate tothe combustor section.

Also disclosed is a method that includes determining, by a controller, aflight phase of an aircraft. The controller determines an operatingparameter of a gas turbine engine of the aircraft, where the gas turbineengine includes a compressor section, a combustor section, and a turbinesection surrounded by an engine case and defining a primary flow path.An active variable cooling flow is adjusted through a bypass ductexternal to the engine case to a secondary flow cavity of the turbinesection based on either or both of the flight phase and the operatingparameter of the gas turbine engine.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where adjusting anactive variable cooling flow includes actuating a cooling air meteringvalve operatively coupled to the bypass duct.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include controlling theactive variable cooling flow through an airflow path within a staticstructure of the gas turbine engine between the bypass duct and thesecondary flow cavity of the turbine section.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include extracting the activevariable cooling flow from a diffuser section proximate to the combustorsection.

A system for an aircraft includes a gas turbine engine having acompressor section, a combustor section, and a turbine sectionsurrounded by an engine case and defining a primary flow path. Thesystem includes a bypass duct external to the engine case and configuredto provide an active variable cooling flow to a secondary flow cavity ofthe turbine section, and at least one cooling air metering valveoperatively coupled to the bypass duct and configured to control theactive variable cooling flow.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the systemincludes a controller configured to actuate the at least one cooling airmetering valve based on either or both of a flight phase and anoperating parameter of the gas turbine engine.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the systemincludes an airflow path within a static structure of the gas turbineengine between the bypass duct and the secondary flow cavity of theturbine section.

A technical effect of systems and methods is achieved by providingactive variable turbine cooling in a gas turbine engine as describedherein.

BRIEF DESCRIPTION OF THE DRAWINGS

The following descriptions should not be considered limiting in any way.With reference to the accompanying drawings, like elements are numberedalike:

FIG. 1 is a schematic illustration of a gas turbine engine in accordancewith an embodiment of the disclosure;

FIG. 2 is a schematic illustration of a cooling system in accordancewith an embodiment of the disclosure;

FIG. 3 is a schematic illustration of a cooling system in accordancewith an embodiment of the disclosure;

FIG. 4 is a schematic illustration of a cooling system in accordancewith an embodiment of the disclosure;

FIG. 5 is a plot of relative internal temperature changes vs. time for avarious flight phases in accordance with an embodiment of thedisclosure; and

FIG. 6 is a flow chart illustrating a method in accordance with anembodiment of the disclosure.

DETAILED DESCRIPTION

A detailed description of one or more embodiments of the disclosedapparatus and method are presented herein by way of exemplification andnot limitation with reference to the Figures.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct, while the compressorsection 24 drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. An engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The engine staticstructure 36 further supports bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis Awhich is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion. It will be appreciated that each of the positions of the fansection 22, compressor section 24, combustor section 26, turbine section28, and fan drive gear system 48 may be varied. For example, gear system48 may be located aft of combustor section 26 or even aft of turbinesection 28, and fan section 22 may be positioned forward or aft of thelocation of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and35,000 ft (10,688 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low correctedfan tip speed” as disclosed herein according to one non-limitingembodiment is less than about 1150 ft/second (350.5 m/sec).

FIG. 2 depicts a cooling system 200 of the gas turbine engine 20according to an embodiment. The cooling system 200 can include a meansfor providing an active variable cooling flow 202 through a bypass duct204 external to an engine case 102 to a secondary flow cavity 206 of theturbine section 28, where the engine case 102 surrounds the compressorsection 24, the combustor section 26, and the turbine section 28. Themeans for providing the active variable cooling flow 202 can include acooling air metering valve 208 that can be actuated between a fullyclosed, partially opened, and fully opened positions. The cooling airmetering valve 208 can be electronically actuated by a controller 290based on either or both of a flight phase and an operating parameter ofthe gas turbine engine 20 as further described herein. In someembodiments, the active variable cooling flow 202 can be routed throughone or more components of the gas turbine engine 20 to reach thesecondary flow cavity 206 internal to the engine case 102. The means forproviding the active variable cooling flow 202 can include an airflowpath 210 within a static structure 212 of the gas turbine engine 20between the bypass duct 204 and the secondary flow cavity 206 of theturbine section 28. The airflow path 210 can be an inner diffuser flowpath configured to deliver a metered supply of cooling air from thecompressor section 24 through a diffuser section 104 of the gas turbineengine 20 proximate to the combustor section 26.

In embodiments, a primary flow path 106 is defined through thecompressor section 24, the combustor section 26, and the turbine section28 as the gas path for compression, combustion, and expansion throughcore components of the gas turbine engine 20. The secondary flow cavity206 provides cooling to reduce the temperature of components downstreamof the combustor 56. Rather than using a mechanically metered flowestablished by flow areas and static holes, embodiments actively controlcooling between a cooler portion of the gas turbine engine 20 and thesecondary flow cavity 206 as dynamically adjusted between a range offlow areas provided by the cooling air metering valve 208. Although asingle instance of the bypass duct 204 and the cooling air meteringvalve 208 is depicted in the example of FIG. 2, a combination ofmetering valves and ducts can be used to control the flow rate of theactive variable cooling flow 202. For example, one or more instances ofthe cooling air metering valve 208 can control the flow rate and volumeof the active variable cooling flow 202. The cooling air metering valve208 can be a continuously variable position valve or may be a discreteposition valve that is modulated (e.g., pulse width modulated) to obtaina partially opened position.

The controller 290 can interface with and control multiple elements ofthe cooling system 200 and the gas turbine engine 20, such as valvestates, flow rates, pressures, temperatures, rotational rates, and thelike. In an embodiment, the controller 290 includes a memory system 292to store instructions that are executed by a processing system 294 ofthe controller 290. The executable instructions may be stored ororganized in any manner and at any level of abstraction, such as inconnection with a controlling and/or monitoring operation of the coolingsystem 200 and/or the gas turbine engine 20. The processing system 294can include one or more processors that can be any type of centralprocessing unit (CPU), including a microprocessor, a digital signalprocessor (DSP), a microcontroller, an application specific integratedcircuit (ASIC), a field programmable gate array (FPGA), or the like.Also, in embodiments, the memory system 292 may include random accessmemory (RAM), read only memory (ROM), or other electronic, optical,magnetic, or any other computer readable medium onto which is storeddata and control algorithms in a non-transitory form.

FIG. 3 depicts a cooling system 300 of the gas turbine engine 20according to an embodiment. The cooling system 300 can include a meansfor providing an active variable cooling flow 302 through a bypass duct304 external to an engine case 102 to a secondary flow cavity 306 of theturbine section 28. The means for providing the active variable coolingflow 302 can include a cooling air metering valve 308 that can beactuated between a fully closed, partially opened, and fully openedpositions. The cooling air metering valve 308 can be electronicallyactuated by the controller 290. In some embodiments, the active variablecooling flow 302 can be routed through one or more components of the gasturbine engine 20 to reach the secondary flow cavity 306 internal to theengine case 102. The secondary flow cavity 306 of the turbine section 28can include a static support structure cooling supply cavity of at leastone stage of vanes 108 of the turbine section 28. Rather than includingcooling holes in a vane support of the vanes 108, the active variablecooling flow 302 can provide cooling between the engine case 102 and thevanes 108. In the example of FIG. 3, the active variable cooling flow302 can be sourced from the diffuser section 104 that can be cooler thanthe primary flow path 106 downstream of the combustor 56.

FIG. 4 depicts a cooling system 400 of the gas turbine engine 20according to an embodiment. The cooling system 400 can include a meansfor providing an active variable cooling flow 402 through a bypass duct404 external to an engine case 102 to a secondary flow cavity 406 of theturbine section 28. The means for providing the active variable coolingflow 302 can include a cooling air metering valve 408 that can beactuated between a fully closed, partially opened, and fully openedpositions. The cooling air metering valve 408 can be electronicallyactuated by the controller 290. In some embodiments, the active variablecooling flow 402 can be routed through one or more components of the gasturbine engine 20 to reach the secondary flow cavity 406 internal to theengine case 102. The means for providing the active variable coolingflow can include a compressor bleed port 410 in the engine case 102proximate to the compressor section 24. The bypass duct 404 can extendexternal to the engine case 102 around the combustor section 26 from thecompressor bleed port 410 to at least one static support structurecooling supply cavity (e.g., secondary flow cavity 406) of the turbinesection 28, such as at least one stage of vanes 108. Although theexamples of FIGS. 2-4 are depicted separately, it will be understoodthat the cooling systems 200, 300, and/or 400 can be combined to providemultiple cooling flow paths for the same instance of the gas turbineengine 20.

FIG. 5 depicts a plot 500 of relative internal temperature changes vs.time for a various flight phases in accordance with an embodiment. Theplot 500 can be established for nominal conditions of the gas turbineengine 20 of FIG. 1 and further augmented based on one or more operatingparameters of the gas turbine engine 20. For example, by monitoringpressures and temperatures of the gas turbine engine 20, it can bedetermined how closely a particular instance of the gas turbine engine20 follows the plot 500, and the use of plot 500 for control actions canbe augmented based on variations observed in one or more operatingparameters of the gas turbine engine 20.

In the example of FIG. 5, an internal temperature of the gas turbineengine 20 is relatively low at idle 502 and climbs rapidly at takeoffpower 504. At climb 506, climb to cruise 508, and subsequent cruisestates 510, 512, 514, the internal temperature of the gas turbine engine20 may gradually reduce as the power demand is reduced and the gasturbine engine 20 is surrounded by cooler ambient air at altitude. Theinternal temperature of the gas turbine engine 20 can be further reducedat decent 516. A temporary spike in the internal temperature of the gasturbine engine 20 may be experienced during a thrust reverse 518operation prior to shut down. A greater flow rate of the active variablecooling flow may be needed at higher temperature flight phases, such asat takeoff power 504, climb 506, and climb to cruise 508. During cruisestates 510, 512, 514, the active variable cooling flow can be reduced.In some embodiments, the active variable cooling flow can besubstantially reduced or partially blocked by the cooling air meteringvalve during decent 516. Upon detecting a temperature rise or inanticipation of thrust reverse 518, the cooling air metering valve canbe opened to increase the active variable cooling flow for a detected oranticipated rapid temperature increase associated with thrust reverse518.

FIG. 6 is a flow chart illustrating a method 600 of actively cooling aturbine section 28 of a gas turbine engine 20 in accordance with anembodiment. The method 600 of FIG. 6 is described in reference to FIGS.1-6 and may be performed with an alternate order and include additionalsteps. The method 600 can be performed, for example, by the systems 200,300, 400 of FIGS. 2-4, a combination thereof, or an alternate bypassconfiguration.

At block 602, controller 290 can determine a flight phase of anaircraft. For example, the flight phase can be determined in referenceto the plot 500. At block 604, controller 290 can determine an operatingparameter of a gas turbine engine 20 of the aircraft. The operatingparameter can be one or more of a pressure and/or temperature within thegas turbine engine 20. At block 606, the controller 290 can adjust anactive variable cooling flow 202, 302, 402 through a bypass duct 204,304, 404 external to the engine case 102 to a secondary flow cavity 206,306, 406 of the turbine section 28 based on either or both of the flightphase and the operating parameter of the gas turbine engine 20.Adjustment of the active variable cooling flow 202, 302, 402 can includeactuating a cooling air metering valve 208, 308, 408 operatively coupledto the bypass duct 204, 304, 404.

The term “about” is intended to include the degree of error associatedwith measurement of the particular quantity based upon the equipmentavailable at the time of filing the application.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the presentdisclosure. As used herein, the singular forms “a”, “an” and “the” areintended to include the plural forms as well, unless the context clearlyindicates otherwise. It will be further understood that the terms“comprises” and/or “comprising,” when used in this specification,specify the presence of stated features, integers, steps, operations,elements, and/or components, but do not preclude the presence oraddition of one or more other features, integers, steps, operations,element components, and/or groups thereof.

While the present disclosure has been described with reference to anexemplary embodiment or embodiments, it will be understood by thoseskilled in the art that various changes may be made and equivalents maybe substituted for elements thereof without departing from the scope ofthe present disclosure. In addition, many modifications may be made toadapt a particular situation or material to the teachings of the presentdisclosure without departing from the essential scope thereof.Therefore, it is intended that the present disclosure not be limited tothe particular embodiment disclosed as the best mode contemplated forcarrying out this present disclosure, but that the present disclosurewill include all embodiments falling within the scope of the claims.

What is claimed is:
 1. A gas turbine engine comprising: a compressorsection; a combustor section; a turbine section operably coupled to thecompressor section, wherein a primary flow path is defined through thecompressor section, the combustor section, and the turbine section; anengine case surrounding the compressor section, the combustor section,and the turbine section; and a means for providing an active variablecooling flow through a bypass duct external to the engine case to asecondary flow cavity of the turbine section.
 2. The gas turbine engineof claim 1, wherein the means for providing the active variable coolingflow comprises a cooling air metering valve.
 3. The gas turbine engineof claim 2, wherein the cooling air metering valve is electronicallyactuated based on either or both of a flight phase and an operatingparameter of the gas turbine engine.
 4. The gas turbine engine of claim1, wherein the means for providing the active variable cooling flowcomprises an airflow path within a static structure of the gas turbineengine between the bypass duct and the secondary flow cavity of theturbine section.
 5. The gas turbine engine of claim 4, wherein theairflow path comprises an inner diffuser flow path configured to delivera metered supply of cooling air from the compressor section through adiffuser section proximate to the combustor section.
 6. The gas turbineengine of claim 1, wherein the secondary flow cavity of the turbinesection comprises a static support structure cooling supply cavity of atleast one stage of vanes of the turbine section.
 7. The gas turbineengine of claim 1, wherein the means for providing the active variablecooling flow comprises a compressor bleed port in the engine caseproximate to the compressor section.
 8. The gas turbine engine of claim7, wherein the bypass duct extends external to the engine case aroundthe combustor section from the compressor bleed port to at least onestatic support structure cooling supply cavity of the turbine section.9. The gas turbine engine of claim 1, wherein the active variablecooling flow is extracted from a diffuser section proximate to thecombustor section.
 10. A method comprising: determining, by acontroller, a flight phase of an aircraft; determining, by thecontroller, an operating parameter of a gas turbine engine of theaircraft, the gas turbine engine comprising a compressor section, acombustor section, and a turbine section surrounded by an engine caseand defining a primary flow path; and adjusting an active variablecooling flow through a bypass duct external to the engine case to asecondary flow cavity of the turbine section based on either or both ofthe flight phase and the operating parameter of the gas turbine engine.11. The method of claim 10, wherein adjusting an active variable coolingflow comprises actuating a cooling air metering valve operativelycoupled to the bypass duct.
 12. The method of claim 10, furthercomprising: controlling the active variable cooling flow through anairflow path within a static structure of the gas turbine engine betweenthe bypass duct and the secondary flow cavity of the turbine section.13. The method of claim 12, wherein the airflow path comprises an innerdiffuser flow path configured to deliver a metered supply of cooling airfrom the compressor section through a diffuser section proximate to thecombustor section.
 14. The method of claim 10, wherein the secondaryflow cavity of the turbine section comprises a static support structurecooling supply cavity of at least one stage of vanes of the turbinesection.
 15. The method of claim 10, wherein the bypass duct extendsexternal to the engine case around the combustor section from acompressor bleed port to at least one static support structure coolingsupply cavity of the turbine section.
 16. The method of claim 10,further comprising: extracting the active variable cooling flow from adiffuser section proximate to the combustor section.
 17. A system for anaircraft, the system comprising: a gas turbine engine comprising acompressor section, a combustor section, and a turbine sectionsurrounded by an engine case and defining a primary flow path; a bypassduct external to the engine case and configured to provide an activevariable cooling flow to a secondary flow cavity of the turbine section;and at least one cooling air metering valve operatively coupled to thebypass duct and configured to control the active variable cooling flow.18. The system of claim 17, further comprising a controller configuredto actuate the at least one cooling air metering valve based on eitheror both of a flight phase and an operating parameter of the gas turbineengine.
 19. The system of claim 17, further comprising an airflow pathwithin a static structure of the gas turbine engine between the bypassduct and the secondary flow cavity of the turbine section.
 20. Thesystem of claim 17, wherein the secondary flow cavity of the turbinesection comprises a static support structure cooling supply cavity of atleast one stage of vanes of the turbine section.